1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine stator vane with film cooling.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
A turbine airfoil—whether a rotor blade or a stator vane—is exposed to different temperatures and pressures from the hot gas stream. Not all areas of the airfoil are adequately cooled, and thus hot spots appear. Hot spots on the airfoils will cause erosion or corrosion damage in which material is lost and holes form. This significantly reduces the useful life of the airfoil. This is especially important for the large frame industrial gas turbine engine used for electric power production because these engines must run for very long periods of time.
A prior art vane is shown in FIG. 1 and includes inner and outer endwalls 12 with the airfoil 11 extending between the endwalls. The vane endwall 12 is cooled using two rows of circular or shaped film cooling holes 13 and 14 located on the upstream side of the endwall. In this cooling design, streamwise and circumferential cooling flow control is difficult to achieve due to the airfoil external hot gas temperature and pressure variations. Film cooling air discharged from the two rows of film cooling holes 13 and 14 have a tendency to migrate from the pressure side toward the vane suction side surface which will induce a mal-distribution of the film cooling flow and endwall metal temperature.